A vertical take-off and landing aircraft

Abstract

The present application relates to the field of aircraft technology, and discloses a vertical take-off and landing aircraft, which comprises a wing-body integrated lifting body fuselage, a front wing and a rear wing, and a stabilizing tail, and is composed of a two-degree-of-freedom propulsion Y3 configuration formed by the front wing, the rear wing and the fuselage, and is composed of two fixed lifting propellers installed below the front wing and in the front part of the gravity center and a two-degree-of-freedom tilting thruster installed at the tail to solve the problem that the existing Y3 configuration cannot simultaneously consider the low-speed hovering mode and the cruising mode; the tandem-lifting body aerodynamic configuration is combined with the new Y3 power configuration to solve the problem that the aircraft is difficult to balance the contradiction between high aerodynamic efficiency and large cargo capacity; and finally, the side aerodynamic optimization of the fuselage, the up-down staggered design of the front and rear wing tip ailerons, and the ring wing 21 design are adopted to solve the problem that the existing configuration is prone to generate large aerodynamic resistance.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aircraft technology, in particular to a vertical take-off and landing aircraft. BACKGROUND

[0002] Currently, the concept aircraft proposed for urban air transportation task application mainly adopts conventional aerodynamic configuration and canard aerodynamic configuration. In order to meet the load demand, the conventional aerodynamic configuration often adopts a large aspect ratio main wing to provide lift, and realizes trim design and flight control through horizontal tail and vertical tail. The aerodynamic configuration is simple to control and mature in performance, but the wing is long and fragile, the structural strength is low, the weight is large, and the space size is large, which restricts the miniaturization and high maneuvering flight of the aircraft; the duck aerodynamic layout moves the horizontal tail of the conventional aerodynamic configuration in front of the main wing, and realizes the pitch trim and control of the aircraft based on the deflection of the canard; compared with the conventional configuration aircraft, the duck aircraft has the advantages of good maneuvering performance, trim resistance and other technologies, but also has the problem of large space size and complex control.

[0003] As can be seen from the above, the Y3 configuration unmanned aerial vehicle currently mainly adopts conventional aerodynamic configuration and canard aerodynamic configuration, but no matter which aerodynamic configuration is adopted, there will be problems in aerodynamics, which can be manifested in two aspects. On the one hand, the aerodynamic configuration is conventional, and it is difficult to balance the contradiction between high aerodynamic efficiency and large load capacity, that is, the aircraft adopts a fuselage to accommodate goods and load, and in order to increase the load capacity and cargo compartment volume, a larger cross-sectional area of the fuselage is often used, which produces a larger waste resistance, further reduces the aerodynamic efficiency, and thus makes it difficult for the aircraft to balance the large capacity and large load performance and high aerodynamic efficiency performance; on the other hand, the Y3 configuration currently lacks aerodynamic optimization, and the aerodynamic efficiency needs to be further improved, that is, mainly reflected in that the Y3 aircraft lacks a transition fairing section of the wing and fuselage, which produces a larger parasitic resistance and interference resistance, that is, the existing aerodynamic configuration produces a larger waste resistance, parasitic resistance and interference resistance, which seriously reduces the aerodynamic efficiency of the aircraft.

[0004] Therefore, in order to realize higher performance indicators to adapt to the demand of urban air transportation task, it is of great significance to research a vertical take-off and landing aircraft which can realize high lift coefficient, high lift ratio, high lift line slope and large available volume aerodynamic configuration. SUMMARY

[0005] The purpose of the present application is to provide a vertical take-off and landing aircraft to solve the problem that the existing vertical take-off and landing aircraft aerodynamic configuration produces a larger waste resistance, parasitic resistance and interference resistance.

[0006] To solve the above technical problems, the application provides a vertical take-off and landing aircraft, which comprises a blended wing body fuselage, a front wing, a rear wing and stabilizer tail wings; the top of the body fuselage is upwardly convex to form an upper convex surface and an upper concave surface, which are arranged in sequence from the nose to the tail, and the curvature of the upper convex surface is larger than that of the upper concave surface; the side of the body fuselage is outwardly convex to form a nose convex part and a tail convex part; the surface of the nose convex part is arc transition with the top surface of the body fuselage and the front end surface of the body fuselage, the highest part of the nose convex part is provided with the front wing, the front wing is provided with a lifting propulsion component, the nose convex part is narrowed from the outer contour line of the nose convex part to the highest part of the nose convex part, and at least part of the nose convex part extends to the tail convex part, and the extension part of the nose convex part is narrowed from the nose to the tail; the surface of the tail convex part is arc transition with the top surface of the body fuselage, the highest part of the tail convex part is provided with the rear wing, and the tail convex part is narrowed from the outer contour line of the tail convex part to the highest part of the tail convex part; the tail of the body fuselage is provided with two stabilizer tail wings, the two stabilizer tail wings are oppositely arranged, the wing tips of the two stabilizer tail wings are upward, and a tail propulsion component is arranged between the two stabilizer tail wings; through the special concave-convex surface design of the top of the body fuselage and the arrangement of the front and rear wings, the aircraft has a larger lifting surface, a better lift-drag ratio, a larger lift curve slope and a large cargo compartment volume, and through the double-convex design of the side of the body fuselage, the drag coefficient of the aircraft can be effectively reduced.

[0007] In one embodiment, in the vertical direction, the installation position height of the front wing is higher than that of the rear wing, the installation angle of the front wing and that of the rear wing are in the same direction, the installation angle of the front wing and that of the rear wing are both less than ninety degrees, and there is an angle difference between the installation angle of the front wing and that of the rear wing; the front wing and the rear wing with the height difference and the angle difference can reduce the effective angle of attack of the rear wing caused by the downward airflow generated by the front wing, thereby causing the lift difference between the front wing and the rear wing, and the lift difference can generate a pitch moment to offset the inherent pitch moment of the wing, and the height difference can modulate the airflow disturbance, so that the pitch moment caused by the aerodynamic interference of the front wing and the rear wing can offset the inherent moment of the wing at a specific wing installation position, and the aircraft can be trimmed.

[0008] In one embodiment, the front wing is installed on the upper part of the nose convex part, and the rear wing is installed on the lower part of the tail convex part; the height difference between the center of gravity of the front wing and that of the rear wing is 30-65 mm, and the angle difference between the installation angle of the front wing and that of the rear wing is 1-2°.

[0009] In one embodiment, the nose bulge comprises a smooth transition mounting head section and an extended tip section; the upper portion of the mounting head section is mounted with the front wing at the junction of the upper portion and the lower portion of the mounting head section; the upper surface of the mounting head section is smoothly transitioned with the surface of the front wing, the lower surface of the mounting head section is connected to the surface of the tail bulge at an obtuse angle, the upper curvature of the mounting head section is within 2° of the lower curvature of the mounting head section; the tip of the extended tip section is tapered in the direction adjacent to the tail bulge; the upper surface of the extended tip section is smoothly transitioned with the surface of the tail bulge, the lower surface of the extended tip section is connected to the surface of the tail bulge at an obtuse angle, the upper curvature of the extended tip section is greater than the lower curvature of the extended tip section; the cooperation of the nose bulge and the front wing on one hand strengthens the structure of the root of the front wing, on the other hand, the smooth transition surface reduces the adverse pressure flow area and the separation area.

[0010] In one embodiment, the tail bulge comprises a smooth transition transition section and a mounting tip section; the head of the transition section is enlarged in the direction adjacent to the nose bulge, the tapered tip of the extended tip section is located on the V-shaped mouth of the transition section, the transition section is connected to the extended tip section; the upper surface of the transition section is smoothly transitioned with the surface of the nose bulge, the lower portion of the transition section extends to the lower portion of the nose bulge, the lower surface of the transition section is connected to the lower surface of the nose bulge at an obtuse angle, the upper curvature of the transition section is greater than the lower curvature of the transition section; the lower portion of the mounting tip section is mounted with the rear wing, the upper surface of the mounting tip section is smoothly transitioned with the surface of the rear wing; the cooperation of the tail bulge and the rear wing on one hand strengthens the structure of the root of the rear wing, on the other hand, achieves good drag reduction effect.

[0011] In one embodiment, the wing tip of the front wing is provided with a front wing tip winglet, and the wing tip of the rear wing is provided with a rear wing tip winglet; the wing tip of the front wing tip winglet is directed to the bottom of the vertical take-off and landing aircraft, and the wing tip of the rear wing tip winglet is directed to the top of the vertical take-off and landing aircraft; the vertical take-off and landing aircraft with such layout design can reduce the induced drag of the aircraft, and the rear wing tip winglet with a larger area can also improve the wing lift and enhance the lateral stability of the aircraft; the downward bending of the front wing tip winglet can guide the wing tip vortex to the lower part of the rear wing, thereby reducing the influence on the rear wing and further reducing the induced drag.

[0012] In one of the embodiments, the out-pitch angle of the front wing tip winglet is 10°, the aspect ratio of the front wing tip winglet is 5%, the out-pitch angle of the rear wing tip winglet is 30°, and the aspect ratio of the rear wing tip winglet is 10%. The front wing tip winglet directly suppresses the wing tip vortex and reduces the induced drag. However, due to the small area and small out-pitch angle, the lift-increasing effect of the front wing tip winglet is weaker than that of the rear wing tip winglet. The large rear wing tip winglet can provide additional yaw correction moment for the aircraft, thereby enhancing the lateral stability of the aircraft.

[0013] In one of the embodiments, the vertical take-off and landing aircraft further comprises a ring wing. The ring wing is installed at the bottom of the front wing, the root of the ring wing is connected to the fuselage, the wing tip of the ring wing is bent towards the top of the vertical take-off and landing aircraft, and the wing tip of the ring wing is connected to the front wing. The wing length of the ring wing is less than or equal to the wing length of the front wing. Compared with the traditional support directly fixed on the wing, the ring wing device effectively reduces the parasitic drag caused by the thrust support and provides part of the lift. At the same time, the ring wing device optimizes the lift and the transmission path of the propeller vibration from the front wing to the fuselage, thereby reducing the overall deformation.

[0014] In one of the embodiments, the relative thickness of the lift body fuselage is at least greater than 11%, and the length of the lift body fuselage is less than 1.5 m.

[0015] In one of the embodiments, the relative thickness of the front wing and the rear wing is at least greater than 11%, the wing chord length of the front wing and the rear wing is 200-400 mm, and the wing span of the front wing and the rear wing is 2-4 m.

[0016] The beneficial effects of the present application are as follows:

[0017] I. Since the fuselage of the vertical take-off and landing aircraft is a lift body fuselage, the top of the lift body fuselage is upwardly convex to form an upper arc convex surface and an upper arc concave surface arranged in sequence from the nose to the tail, and the front wing and the rear wing are arranged in series to form a lift body configuration. When flying, the fuselage adopts a wide and wing-like design, and the smooth transition of the fuselage and the wing design can greatly reduce the drag of the aircraft fuselage and generate part of the lift. The aerodynamic configuration of the series-lift body formed by the aircraft can provide a larger lift surface, a higher lift-drag ratio, a larger lift curve slope, and a large cargo compartment volume, so as to realize the heavy load, long endurance and high maneuverability of the vertical take-off and landing aircraft, and better meet the needs of the aircraft to shuttle between buildings and maintain low-speed hovering in urban air traffic.

[0018] II. The side of the lifting body fuselage is enlarged and thickened by the design of the nose protrusion and the tail protrusion, the nose protrusion can strengthen the stability of the front wing root structure, similarly, the tail protrusion can strengthen the stability of the rear wing root structure, and the nose protrusion and the tail protrusion are smoothly transitioned, which can effectively reduce the adverse pressure flow area and the separation area, significantly reduce the wing-body interference drag of the aircraft, and enhance the cruising efficiency of the aircraft.

[0019] In summary, the tandem-lifting body type configuration formed by the special designed lifting body fuselage, the front wing and the rear wing can realize the aerodynamic configuration aircraft with high lift coefficient, high lift ratio, high lift line slope and large available volume. BRIEF DESCRIPTION OF DRAWINGS

[0020] In order to more clearly illustrate the technical solutions of the present application, the drawings needed in the embodiments will be briefly introduced as follows. Obviously, the drawings in the following description are only some embodiments of the present application, and other drawings can also be obtained by those skilled in the art without creative labor.

[0021] Figure 1 is a top view schematic diagram of the overall aircraft structure provided by the preferred embodiment of the present application;

[0022] Figure 2 is a bottom view schematic diagram of the overall aircraft structure provided by the preferred embodiment of the present application;

[0023] Figure 3 is a side view schematic diagram of the overall aircraft structure provided by the preferred embodiment of the present application;

[0024] Figure 4 is a top view schematic diagram of the overall aircraft structure provided by the preferred embodiment of the present application;

[0025] Figure 5 is a local large-scale diagram of the nose protrusion of the side of the lifting body fuselage provided by the preferred embodiment of the present application;

[0026] Figure 6 is a local large-scale diagram of the tail protrusion of the side of the lifting body fuselage provided by the preferred embodiment of the present application;

[0027] Figure 7 is a fairing curve design and cloud diagram of the overall aircraft provided by the preferred embodiment of the present application;

[0028] Figure 8 is a parameterization diagram of the relative position of the front wing and the rear wing of the overall aircraft provided by the preferred embodiment of the present application;

[0029] Figure 9 is a typical tandem wing pitch moment variation and lift-drag characteristic diagram provided by the present application;

[0030] Figure 10 is a whole aircraft wing tip device cloud diagram provided by the preferred embodiment of the present application;

[0031] Figure 11 is a hovering power consumption and tail motor hovering lift diagram provided by the preferred embodiment of the present application;

[0032] Figure 12 is a local structure Fluent calculation verification diagram provided by the preferred embodiment of the present application;

[0033] Figure 13 is a power consumption diagram provided by the preferred embodiment of the present application.

[0034] The reference signs are as follows:

[0035] 1, lifting body fuselage; 10, upper arc convex surface; 11, upper arc concave surface; 12, nose convex; 120, mounting head section; 121, extended tail section; 13, tail convex; 130, transition middle section; 131, mounting tail section;

[0036] 2, front wing; 20, front winglet; 21, ring wing;

[0037] 3, rear wing; 30, rear winglet;

[0038] 4, stabilizer; 5, lifting propulsion component; 6, tail propulsion component. DETAILED DESCRIPTION

[0039] The technical solutions in the embodiments of the present application will be described clearly and completely below with reference to the drawings in the embodiments of the present application.

[0040] Since the eVTOL aircraft needs to realize the rotor mode vertical take-off and landing and the fixed wing mode horizontal cruising, it generally has a propulsion system in two flight modes, or a tilt rotor propulsion system. According to the arrangement, function and form of the propeller, the current eVTOL aircraft configuration types can be roughly divided into: tail seat type vertical take-off and landing aircraft, composite type vertical take-off and landing aircraft, full tilt thrust type vertical take-off and landing aircraft, and partial tilt thrust type vertical take-off and landing aircraft. The Y3 configuration concerned in the present application belongs to the partial tilt thrust type vertical take-off and landing aircraft. The unmanned aerial vehicle of the configuration is composed of two tilt propellers installed below the front wing near the center of gravity and a tail fixed propeller. However, the existing Y3 configuration vertical take-off and landing aircraft has several problems, which are as follows.

[0041] First, the existing Y3 configuration eVTOL cannot balance the high power efficiency design of the low-speed hovering mode and the cruising mode. Specifically, the two tilting propellers in front of the traditional Y3 configuration unmanned aerial vehicle are close to the center of gravity, bearing most of the hovering lift, and need to be adapted to the high power efficiency design of the hovering mode. However, the propeller also needs to provide all the propulsion of the cruising mode, so it needs to be adapted to the high power efficiency design of the cruising mode. The low-speed hovering mode has a small incoming flow speed, and the high power efficiency design requires large blade size and low motor speed. The cruising mode has a large incoming flow speed, and the high power efficiency design requires small blade size and high motor speed. Because the high power efficiency design in the two modes requires opposite blade size and motor speed, it is difficult to balance the two design requirements, resulting in that the traditional Y3 configuration has low comprehensive power efficiency in the two modes.

[0042] Second, the existing conventional aerodynamic layout eVTOL aircraft for urban air traffic cannot balance the contradiction between high aerodynamic efficiency and large cargo capacity. Specifically, to improve the cargo capacity and cargo compartment volume, the existing eVTOL aircraft often uses a larger cross-sectional area of the fuselage, which reduces the lift-drag ratio and aerodynamic efficiency of the aircraft.

[0043] Third, the aerodynamic design of the existing Y3 aircraft needs to be optimized in detail. Specifically, the existing Y3 aerodynamic shape lacks a transition fairing between the wing and the fuselage, resulting in large parasitic drag and interference drag. The existing Y3 configuration aircraft lacks winglets, resulting in large induced drag, which reduces the aerodynamic efficiency of the aircraft and affects the endurance. The front propeller of the existing Y3 is basically connected to the front wing surface through a support, which affects the flow of the wing surface and generates large parasitic drag.

[0044] Therefore, the present application solves the above problems through three major directions of design. First, a two-degree-of-freedom propulsion Y3 configuration is proposed, which is composed of two fixed lift propellers installed below the front wing and in the front part of the center of gravity, and a two-degree-of-freedom tilting propeller installed at the tail, to solve the problem that the existing Y3 configuration eVTOL cannot balance the low-speed hovering mode and the cruising mode. Second, the combination of tandem-lift body aerodynamic configuration and new Y3 power configuration is adopted to solve the problem that the aircraft cannot balance the contradiction between high aerodynamic efficiency and large cargo capacity. Third, the side aerodynamic optimization of the fuselage, the up-down staggered design of the front and rear winglets, and the ring wing 21 design are used to solve the problem that the existing configuration is prone to generate large aerodynamic drag.

[0045] In the embodiments of the present application, please refer to Figures 1 to 6The vertical take-off and landing aircraft provided by the application comprises a lifting body fuselage 1, a front wing 2, a rear wing 3 and a stabilizer tail 4; two front wings 2 are symmetrically arranged on the two sides of the lifting body fuselage 1, two rear wings 3 are also symmetrically arranged on the two sides of the lifting body fuselage 1, the front wing 2 and the rear wing 3 are connected with the lifting body fuselage 1 in a wing-body integrated manner, two stabilizer tails 4 are symmetrically arranged on the tail of the lifting body fuselage 1, the wing tips of the stabilizer tails 4 are upward, and the front wing 2 is provided with a lifting propulsion component 5, and a tail propulsion component 6 is arranged between the two stabilizer tails 4 to provide the aircraft with lifting force. After the arrangement is adopted, the front wing 2 and the rear wing 3 generate lifting force together with the lifting body fuselage 1, compared with the same weight level eVTOL aircraft, the configuration is compact, the space size is greatly reduced, and flexible take-off and landing in a small space are facilitated.

[0046] It should be noted that the wing-body integrated wing refers to that the wing and the fuselage are designed as a whole, the planar shape and the cross-sectional shape of the two are completely integrated as a whole, the layout of the fuselage and the wing is smoothly transitioned without obvious boundary, the wing integration can make the aircraft obtain better aerodynamic performance, and the wing integration body has the advantages of light structure, large volume and small resistance.

[0047] In the embodiment of the application, the lifting body fuselage 1 is described above, please refer to Figure 3 and Figure 4 The lifting body fuselage 1 adopts a wide fuselage design as a whole; the top of the lifting body fuselage 1 is upwardly convex to form an upper arc convex surface 10 and an upper arc concave surface 11, the upper arc convex surface 10 and the upper arc concave surface 11 are arranged in sequence from the nose to the tail, and the curvature of the upper arc convex surface 10 is greater than that of the upper arc concave surface 11; the side of the lifting body fuselage 1 is outwardly convex to form a nose convex portion 12 for transition installation of the front wing 2 and a tail convex portion 13 for transition installation of the rear wing 3. After the arrangement is adopted, the wide lifting body design of the lifting body fuselage 1 greatly reduces the interference resistance to realize high lift-drag ratio, efficient flight, and provides sufficient volume space for the cargo compartment load. Moreover, the side of the lifting body fuselage 1 is provided with the nose convex portion 12 and the tail convex portion 13 formed by special design, so that the wing-body interference resistance of the aircraft is significantly reduced, the cruising efficiency of the aircraft is enhanced, and the aircraft structure design and the fuselage cargo compartment design are facilitated.

[0048] As shown in Figure 2 , the bottom of the lifting body fuselage 1 is provided with an arc-shaped frame and a support frame, the arc-shaped frame is arranged at the front end of the lifting body fuselage 1, and the support frame is arranged at the tail end of the lifting body fuselage 1.

[0049] In the embodiment of the application, as shown in Figure 4As shown, the side of the lifting body fuselage 1 is mainly composed of the nose protrusion 12 and the tail protrusion 13, and the side of the lifting body fuselage 1 is crucial for the proper design of the side of the lifting body fuselage 1 (hereinafter referred to as the transition section) of the blended wing lifting body aircraft, therefore, the design method of the side of the lifting body fuselage 1 will be given below, which specifically includes the following steps:

[0050] 1) The intersection of the two lofting surfaces is obtained by using a double quadratic curve fitting lofting, realizing smooth transition of the overall geometry without obvious interference points;

[0051] The lofting formula of the quadratic surface is as follows:

[0052] Transition section curve one: P(u) = Au 2 + Bu + C

[0053] Transition section curve two: Q(v) = Dv 2 + Ev + F

[0054] Transition section curve one ( Figure 7 curve one), curve two ( Figure 7 curve two), curve one represents the profile curve of the side of the lifting body fuselage, and curve two represents the profile curve of the wing root airfoil. Wherein u, v represent parameters, (u, P(u)) represents a point on the transition curve one, and P(u) traverses every point on the transition curve one as the parameter u changes. Q(v) is the same.

[0055] Define the lofting parameter t∈[0,1]. Then through interpolation, the parameter t is interpolated between the two curves P(u) and Q(v):

[0056] R(t) = P(u)*(1-t) + Q(v)*t

[0057] Where R(t) represents a point on the lofting surface generated, which gradually transitions from one curve to another as t changes in [0,1], thereby generating the entire surface.

[0058] The above-mentioned quadratic surface lofting method is commonly used in the design of wing-body transition sections, but the existing transition fairing design is from one fuselage to a pair of wings. Based on the consideration of the tandem wing aerodynamic configuration, a transition fairing design method from one fuselage to two pairs of wings is needed.

[0059] Therefore, this patent uses the quadratic surface lofting method to design the transition fairing of the front and rear wings 3 respectively, and then performs Boolean addition (set union) to generate a transition fairing from one fuselage to two pairs of wings. This method is called double quadratic surface lofting.

[0060] 2) The surface pressure coefficient distribution of the side of the lifting body fuselage 1 (transition surface) obtained by the CFD (Computational Fluid Dynamics) method, identify the flow separation, adverse pressure gradient and flow state;

[0061] 3) The guide line of the front wing 2 root to the wing belly, the front wing 2 leading edge to the nose, the rear wing 3 tail to the tail of the fuselage (the guide line of the front wing 2 root to the wing belly) Figure 7 The guide line finally realizes the reasonable arrangement of the curve transition, reduces the adverse pressure flow area and the separation area. Figure 7

[0062] The purpose of the guide line is to reduce the adverse pressure flow area and the separation area, and the guide line is shown in Figure 7 The guide line separates the originally smooth transition section.

[0063] In the embodiment of the present application, the spanwise length of the side of the lifting body fuselage 1 (i.e. the transition section) obtained by the design method of the side of the lifting body fuselage 1 is 90mm, and the width of the entire fuselage is 580mm.

[0064] In some embodiments of the present application, as shown in Figure 7 The structure of the nose bulge 12 generated, as shown in Figure 1 , Figure 4 and Figure 5 The surface of the nose bulge 12 and the top surface of the lifting body fuselage 1, the front end surface of the lifting body fuselage 1 are all circular arc transitions, the nose bulge 12 is installed on the highest part of the nose bulge 12, the front wing 2 is provided with a lifting propulsion component 5, the nose bulge 12 narrows from the outer contour line of the nose bulge 12 to the highest part of the nose bulge 12, and at least part of the nose bulge 12 extends to the tail bulge 13, the extension part of the nose bulge 12 narrows from the nose to the tail direction, after adopting this setting mode, on the one hand, the nose bulge 12 can smoothly transition the front wing 2 and the fuselage, reasonably arrange the curve transition of the two, reduce the adverse pressure flow area and the separation area, on the other hand, the nose bulge 12 can increase the thickness of the connection section of the wing body, thereby realizing the strengthening of the structure of the front wing 2 root.

[0065] ​Specifically, the nose protrusion 12 comprises a smooth transition mounting head section 120 and an extended end section 121; the upper portion of the mounting head section 120 is connected to the lower portion of the mounting head section 120 at the junction, the upper surface of the mounting head section 120 is smoothly transitioned with the surface of the front wing 2, the lower surface of the mounting head section 120 is connected to the surface of the tail protrusion 13 in an obtuse angle manner, the curvature difference between the upper curvature of the mounting head section 120 and the lower curvature of the mounting head section 120 is 2°; the tip of the extended end section 121 is tapered in the direction adjacent to the tail protrusion 13; the upper surface of the extended end section 121 is smoothly transitioned with the surface of the tail protrusion 13, the lower surface of the extended end section 121 is connected to the surface of the tail protrusion 13 in an obtuse angle manner, the upper curvature of the extended end section 121 is greater than the lower curvature of the extended end section 121 (the curvature difference is also 2°), after adopting this arrangement, the nose protrusion 12 can achieve smooth transition of the wing-body and strengthen the structure of the root of the front wing 2.

[0066] Specifically, the obtuse angle connection between the lower surface of the mounting head section 120 and the surface of the tail protrusion 13 means that the included angle between the lower surface of the mounting head section 120 and the surface of the tail protrusion 13 is an obtuse angle, and the lower portion of the mounting head section 120 and the surface of the tail protrusion 13 are fixedly connected.

[0067] It is particularly pointed out that the curvature difference in this example can be up and down within the error allowable range.

[0068] In some embodiments of the present application, as Figure 7 The structure of the generated tail protrusion 13 is as shown in Figure 1 , Figure 4 and Figure 6 , the surface of the tail protrusion 13 is arc transitioned with the top surface of the lifting body fuselage 1, the highest protrusion of the tail protrusion 13 is installed with the rear wing 3, the tail protrusion 13 is tapered from the outer contour line of the tail protrusion 13 to the highest protrusion of the tail protrusion 13, after adopting this arrangement, its effect is similar to that of the nose protrusion 12, which can also achieve smooth transition of the rear wing 3 and the fuselage, reasonably arrange the curve transition of the two, reduce the adverse pressure flow area and the separation area, on the other hand, the nose protrusion 12 can increase the thickness of the connecting section of the wing-body, thereby achieving the strengthening of the root structure of the rear wing 3.

[0069] Specifically, the tail protrusion 13 includes a smooth transition transition middle section 130 and an installation end section 131; adjacent to the direction of the nose protrusion 12, the leading end of the transition middle section 130 is an enlarged port, and the tapered tip of the extended end section is arranged on the V-shaped port of the transition middle section 130, and the transition middle section 130 is matched and connected with the extended end section 121; the upper surface of the transition middle section 130 is smoothly transitioned with the surface of the nose protrusion 12, the lower part of the transition middle section 130 extends to the lower part of the nose protrusion 12, the lower surface of the transition middle section 130 is at an obtuse angle with the lower surface of the nose protrusion 12, the upper arc of the transition middle section 130 is greater than the lower arc of the transition middle section 130; the lower part of the installation end section 131 is installed with the rear wing 3, and the upper surface of the installation end section 131 is smoothly transitioned with the surface of the rear wing 3.

[0070] In the embodiments of the present application, the front wing 2 and the rear wing 3 are as shown in Figure 1 and Figure 2 The front wing 2 and the rear wing 3 are in the vertical direction, the installation position height of the front wing 2 is higher than the installation position height of the rear wing 3, the installation angle of the front wing 2 and the installation angle of the rear wing 3 are in the same direction, that is, the leading edge of the front wing 2 is installed in the nose direction, and the leading edge of the rear wing 3 is also installed in the nose direction, the installation angle of the front wing 2 and the installation angle of the rear wing 3 are both less than ninety degrees, and there is an angle difference between the installation angle of the front wing 2 and the installation angle of the rear wing 3. After adopting this setting mode, higher lift coefficient, larger lift line slope and better cruising performance can be achieved.

[0071] However, due to the complex, nonlinear, and strong coupling aerodynamic interference between the front wing 2 and the rear wing 3, that is, the downwash airflow of the front wing 2 affects the aerodynamic characteristics of the rear wing 3, and the induced flow field of the rear wing 3 affects the aerodynamic characteristics of the front wing 2; this influence causes the change of the lift distribution of the front wing 2 and the rear wing 3 of the aircraft, and further affects the pitch moment balance and focus position of the aircraft.

[0072] Therefore, in order to realize the effective balance design of the aircraft with this configuration, the design method of the relative position of the front wing 2 and the rear wing 3 is given, which specifically includes the following steps:

[0073] 1) Establish an aircraft basic aerodynamic geometric parameterization model;

[0074] Among them, for the tandem-lift body configuration, the relative installation position parameters of the front wing 2 and the rear wing 3 are as shown in Figure 8The relative position of the front and rear wings 3 in the vertical plane before and after quantization is the vertical relative distance (gap) and the horizontal relative distance (X) based on the body coordinate system. The vertical relative distance is the projection of the relative distance between the leading edge points of the wing airfoils at the root of the two wings on the z-axis (vertical direction) of the aircraft body coordinate system, and the value is positive when the front wing 2 is placed below the rear wing 3. The horizontal relative distance is the projection of the relative distance between the aerodynamic centers of the wing airfoils at the root of the front and rear wings 3 on the x-axis (axial direction) of the body coordinate system.

[0075] 2) Select the highest lift-drag ratio angle of attack of the fuselage airfoil based on the aerodynamic geometric model, determine the cruise angle of attack of the aircraft, and subtract the cruise angle of attack from the optimal lift-drag ratio angle of attack of the wing airfoil to obtain the initial installation angle of the front and rear wings 3, so as to ensure that the front and rear wings 3 work near the optimal lift-drag ratio angle of attack in the cruise state;

[0076] Wherein, the relative angle of the front and rear wings 3 is the decangle, and the value is positive when the installation angle of the front wing 2 is greater than that of the rear wing 3. With the increase of the decangle, the zero-lift pitch moment of the aircraft decreases rapidly and linearly, which is consistent with the theoretical explanation. With the gradual increase of the installation angle of the front wing 2 compared with the rear wing 3, the lift difference between the front wing 2 and the rear wing 3 becomes more and more obvious. Since the lift coefficient of the wing changes almost linearly with the angle of attack in the non-stall working section, and the relative installation angle of the front and rear wings 3 does not change during the change of the angle of attack of the aircraft, the difference between the two is almost constant and approximately equal to the decangle. Therefore, the lift difference changes approximately as a constant with the change of the angle of attack of the aircraft, resulting in the pitch moment that offsets the inherent pitching moment of the wing. With the gradual increase of the decangle, the moment caused by the lift difference gradually increases, so the pitch moment of the whole machine changes approximately linearly.

[0077] 3) Based on the CFD method to solve the lift-drag characteristics of the aircraft, the focus position and the pitch moment, and adjust the gap so that the zero-lift pitch moment of the aircraft is close to 0;

[0078] It is worth noting that with the change of the gap, the pitch moment of the aircraft also shows a similar phenomenon to the change of the decangle. With the decrease of the gap, the zero-lift pitch moment of the aircraft decreases nonlinearly, and the pitch moment is zero at a certain value, realizing the pitch trim. This is due to the nonlinear aerodynamic interference of the front and rear wings 3; compared with the lift-drag characteristics of the front and rear wings 3 themselves, the downwash flow generated by the front wing 2 reduces the effective angle of attack of the rear wing 3, thereby causing the lift difference between the front and rear wings 3; the lift difference produces a pitch moment that offsets the inherent pitching moment of the wing; and the gap has a certain modulation on the airflow disturbance, and at a certain wing installation position, the pitch moment caused by the aerodynamic interference of the front and rear wings 3 can offset the inherent moment of the wing, realizing the trim of the aircraft.

[0079] It should be noted that the relationship between the zero-lift pitching moment of the aircraft and the gap makes it possible to adjust the gap to generate a pitching moment by using the airflow interference between the front and rear wings 3, and to offset part of the inherent nose-down moment generated by the wings, thus achieving gap-based pitching trim.

[0080] 4) Fine-tune the decangle based on the above design model to make the aircraft meet the set flight condition for complete pitching moment trim.

[0081] Compared with the method of not adjusting the gap, the method of adjusting only the decangle to achieve pitching trim greatly reduces the wing difference angle of the aircraft; the wing installation angle of the aircraft has a huge impact on the lift-drag characteristics, and adjusting the wing difference angle means adjusting the wing installation angle to deviate from its optimal design angle; therefore, the tandem aircraft based on gap trimming has significantly better lift-drag characteristics, with a significantly improved lift coefficient and lift line slope, and a smaller cruise attack angle, as shown in Figure 9 .

[0082] In some embodiments of the present application, the relative positions of the front wing 2 and the rear wing 3 are determined by using the above design method, as shown in Figure 8 In this embodiment, the height difference (gap) between the center of gravity of the front wing 2 and the center of gravity of the rear wing 3 is 65 mm, the horizontal distance difference (X) between the center of gravity of the front wing 2 and the center of gravity of the rear wing 3 is 550 mm, the installation angle of the front wing 2 is 2°, the installation angle of the rear wing 3 is 0°, and the angle difference between the installation angles of the two wings is 2°. After using this setting method, the lift center position of the entire aircraft is 0.5886 m from the nose.

[0083] In some other embodiments of the present application, the relative positions of the front wing 2 and the rear wing 3 are also determined by using the above design method, in which the height difference (gap) between the center of gravity of the front wing 2 and the center of gravity of the rear wing 3 is 30 mm, the horizontal distance difference (X) between the center of gravity of the front wing 2 and the center of gravity of the rear wing 3 is 550 mm, the installation angle of the front wing 2 is 2°, the installation angle of the rear wing 3 is 1°, and the angle difference between the installation angles of the two wings is 1°. After using this setting method, the lift center position of the entire aircraft is 0.5819 m from the nose.

[0084] The gap-based pitching trim method uses the unique front-rear wing aerodynamic interference phenomenon of the tandem aircraft to trim the aircraft, and compared with the method of not adjusting the gap, the installation height and installation angle of the front wing 2 and the rear wing 3 are more reasonable, so that the entire aircraft can achieve a higher lift coefficient, a larger lift line slope, and better cruise performance.

[0085] For the forewing 2, the relative thickness of the forewing 2 is at least greater than 11%, the wing chord length of the forewing 2 is 300mm, and the wingspan of the forewing 2 is 1.8m.

[0086] In some embodiments of this application, in order to reduce the parasitic drag caused by the lifting support, such as Figure 1 and Figure 2 As shown, a ring wing 21 is installed at the bottom of the front wing 2. The root of the ring wing 21 is connected to the fuselage, and the wingtip of the ring wing 21 bends towards the top of the vertical take-off and landing aircraft. The wingtip of the ring wing 21 is connected to the bottom of the front wing 2. A lift thruster is installed at the bottom of the ring wing 21. The wingspan of the ring wing 21 is less than that of the front wing. By adopting this configuration, the ring wing 21 replaces the existing technology of supporting the lift thruster with a bracket, effectively reducing the parasitic drag caused by the traditional lift bracket, and realizing the integrated optimization design of the lift thruster and the aerodynamic model of the aircraft.

[0087] Specifically, the Ring Wing 21 airfoil is a small-camber RG15A airfoil with an overall installation angle of 0° relative to the fuselage axis, a wingspan of 335mm, a root leading edge distance of 190mm from the fuselage leading edge, a root chord length of 100mm, and a vertical spacing of 85mm between the upper and lower wings. This configuration reduces the impact of the Ring Wing 21 on the aerodynamic characteristics of the aircraft and minimizes the lift-to-drag ratio loss caused by the Ring Wing 21.

[0088] In some embodiments of this application, in order to reduce the induced drag of the aircraft and further improve the lift-drag characteristics, such as Figure 2 and Figure 10 As shown, the front wing 2 has a front wingtip winglet at its wingtip. The wingtip of the front wingtip winglet 20 faces the bottom of the vertical take-off and landing aircraft. With this configuration, the front wingtip winglet 20 is a downward-curved wingtip winglet, which can directly suppress wingtip vortices and reduce induced drag.

[0089] Specifically, the leading winglet uses a small, downward-curved winglet with a cant angle of 10° and a span ratio of 5%. Viewed from the side, the winglet has a 48° angle at the front, a wingtip width of 160mm, and a downward curve of 100mm. Viewed from above, the wingtip protrudes 35mm outward in the span direction. With this configuration, since the leading wing 2 is positioned in front of the aerodynamic center of the aircraft, if a larger winglet were used at the tip of the leading wing 2, it would reduce the effective lateral stabilizing area of ​​the aircraft, leading to lateral instability. Therefore, the leading wing 2 should use a smaller winglet. Furthermore, according to CFD calculations, the smaller area of ​​the leading wing 2 results in greater lift, leading to a higher actual wing loading compared to the fuselage and rear wing 3, and a smaller aspect ratio. Therefore, it is foreseeable that it will generate a more pronounced wingtip vortex. This wingtip vortex interacts with the rear wing 3, producing significant induced drag.

[0090] For the rear wing 3, the relative thickness of the front wing 2 is at least greater than 11%, the wing chord length of the front wing 2 is 300mm, and the wingspan of the front wing 2 is 3m.

[0091] In some embodiments of this application, also in order to reduce the induced drag of the aircraft, such as Figure 2 and Figure 10 As shown, the rear wing 3 has a rear wingtip 30 at its wingtip. The wingtip of the rear wingtip faces the top of the vertical take-off and landing aircraft. With this configuration, the rear wingtip is an upward-curved wingtip, which can reduce wingtip vortices and reduce the induced drag of the aircraft. On the other hand, the wingtip with a larger outward cant angle interacts with the vortex, which can improve the lift of the wing.

[0092] Specifically, the rear winglets are large, upward-curved winglets with a cant angle of 30° and a span-to-length ratio of 10%. With this configuration, the larger rear winglets provide additional yaw-correcting moment for the aircraft, thereby enhancing its lateral stability.

[0093] By employing a smaller, downward-curved wingtip on the front wingtip and a larger, upward-curved wingtip on the rear wingtip, the induced drag of the aircraft is reduced. Simultaneously, the rear wingtip, with its larger area, can also increase wing lift and enhance lateral stability. The downward curve of the front wingtip directs the generated wingtip vortex to below the rear wing, reducing its impact and thus decreasing induced drag. This is achieved by obtaining the pressure coefficient distribution and streamline diagram near the device using a CFD method. Figure 10 It can be seen that the wingtip vortex suppression is obvious, the pressure coefficients on the upper and lower surfaces of the winglet are smoothly transitioned, and the induced flow field caused by the wingtip of the forewing 2 does not have a significant impact on the flow of the rear wing 3.

[0094] In the embodiments of this application, please refer to Figure 1 Regarding the aforementioned lift propulsion component 5 and tail propulsion component 6, this scheme adopts a compound propulsion configuration and a tilt propulsion configuration to form a two-degree-of-freedom propulsion configuration. The two lift propulsion components 5 are respectively located at the bottom of the two forewings 2. The lift propulsion components 5 are used to generate the main lift of the aircraft in the low-speed hovering mode. The tail propulsion component 6 is located below the tail of the lifting body fuselage 1 and is arranged between the two stabilizer fins 4. The tail propulsion component 6 is used to generate a small portion of the lift in the low-speed hovering mode and the propulsion force of the aircraft in the cruise mode. Since the tail propulsion component 6 can achieve two-degree-of-freedom tilting towards the bottom of the fuselage and towards the side of the fuselage, it can achieve better longitudinal and directional deflection of the propeller, thereby realizing the vertical take-off and landing of the aircraft and the yaw control of the aircraft in the low-speed hovering and transition modes.

[0095] For the tail propulsion component 6, such as Figure 1The tail propulsion component 6 is connected to a tail fixing frame, which is located on the top surface of the tail of the lifting body fuselage 1. The arrangement axis of the tail fixing frame is in the same direction as the axis of the lifting body fuselage 1. The tail fixing frame allows the tail propulsion component 6 to be arranged below the tail of the lifting body fuselage 1.

[0096] It should be noted that, compared to other propulsion configurations of eVTOL aircraft, the main technical advantages of adopting a two-degree-of-freedom propulsion configuration include:

[0097] 1) Large-size, low-speed lift thrusters optimized for hovering mode and small-size, high-pitch tilt thrusters optimized for cruise mode are respectively used to generate the main lift force for low-speed hovering mode and the thrust force for cruise mode, thus taking into account the high propulsion efficiency of both low-speed hovering mode and cruise mode.

[0098] 2) By using yaw-deflectable thrusters to control the aircraft's low-speed hovering and transition mode yaw, the aircraft's yaw control margin is greatly improved, and the aircraft's control reliability is enhanced.

[0099] 3) During flight, fewer thrusters are shut down, and there are no complex structural supports. Furthermore, the two-degree-of-freedom tail tilt thruster can participate in the aircraft's cruise control, thus significantly reducing the aircraft's dead weight. These technological advantages enable eVTOL aircraft using this configuration to better adapt to the mission requirements of urban air transport, including long hovering and cruise durations, high maneuverability, and large payloads.

[0100] In this embodiment, the specific parameters and selection of the lift thruster and tail thruster are based on the 10kg-class eVTOL aircraft for logistics transportation that this study is aimed at, and the specific selection is as follows based on the takeoff weight and class.

[0101] The motor of the lifting body thruster is preferably U8IIKV100, the propeller is preferably 27*8.8, the maximum thrust is preferably 8.4kg, the rated thrust is preferably 2.5~3.5kg, and the rated force efficiency is preferably 10.5g / w.

[0102] The motor for the tail thruster is preferably an XM5060KV220, the propeller is preferably 16cm*8cm, the maximum thrust is preferably 9.5kg, the rated thrust is preferably 2.5~4kg, and the rated force efficiency is preferably 7.2g / w.

[0103] To rationally allocate lift between the lift thrusters and tail thrusters, the following section will further analyze the hovering lift distribution problem of this scheme. We define e1(F1) as the lift thruster efficiency of the forewing 2 in hovering mode, its magnitude varying with its output thrust F1; similarly, we define the tail thruster efficiency e2(F2); and the hovering mode power consumption W... h :

[0104]

[0105] In a hovering state, the lift from the thrusters equals the weight, that is:

[0106] G To =2F1+F2

[0107] Substituting these values, we obtain the relationship between the hovering mode power consumption and the ratio of tail motor thrust to takeoff weight:

[0108]

[0109] Based on the motor-related data, e1(F1) and e2(F2) are plotted, as follows: Figure 11 As shown in the left figure, further, draw W. h The relationship with p2 is as follows Figure 11 As shown in the right figure, it can be seen that as the lift generated by the tail thruster gradually increases, the hovering power consumption of the aircraft first decreases slowly and then gradually increases, which is consistent with the previous physical analysis results. Furthermore, the hovering power consumption is almost constant and minimal when the tail thruster allocates a lift percentage between 10% and 30%. Taking into account the aircraft's control performance, the lift distribution scheme is set as 30% load for the tail thruster and 35% load for each individual lift thruster. Based on this distribution scheme, the specific installation positions of the tail thruster and lift thruster are determined.

[0110] Typical tilt-and-launch (VTOL) aircraft use thrusters suitable for cruise flight mode to provide lift for hovering mode, and their hovering power consumption includes:

[0111]

[0112] Its power consumption is plotted as follows Figures 2-6 As shown by the dashed line, the tilt configuration of the two-degree-of-freedom tail thruster in this scheme can achieve a significant reduction in hovering power consumption. The current design scheme can reduce hovering power consumption by 40%, which can significantly extend the hovering endurance of the aircraft.

[0113] By adopting the two-degree-of-freedom propulsion configuration in this embodiment, on the one hand, hovering power consumption can be significantly reduced, the propulsion efficiency of the aircraft in low-speed hovering mode can be improved, and a longer hovering flight time can be achieved, which can meet the needs of urban air traffic logistics transportation tasks for long-term hovering and multiple take-offs and landings; on the other hand, hovering and transition mode yaw control based on tail thruster yaw can significantly improve the handling stability and control reliability of the aircraft in the corresponding flight state, which can meet the requirements of high maneuverability, high reliability and high wind resistance missions.

[0114] In this embodiment, the flight control system of the vertical takeoff and landing aircraft mainly includes hovering mode control logic and control logic for control surface deflection.

[0115] For the hovering mode control logic, in the hovering mode of the aircraft, lift is generated jointly by the lift thruster fixed below the forewing wing 2 and the tail two-degree-of-freedom tilt thruster. Among them, the lift thruster is closer to the center of gravity and provides 70% of the hovering lift of the aircraft, while the tail thruster provides the remaining lift and deflects at a small angle to one side to generate a thrust component, which in turn generates a yaw torque to counteract the angular momentum anti-torque caused by its own rotation.

[0116] When the aircraft needs to perform yaw control, the lift thrusters of the front wing 2 are used to maintain lateral stability, and the tail thrusters are used to generate yaw torque. For example, if the aircraft needs to be controlled to yaw to the left, the tail thrusters of the aircraft will deflect laterally to the yaw side, generating yaw torque to turn to the left, thereby achieving yaw control of the aircraft. The lift thrusters of the aircraft will adjust their speed to maintain the lateral and longitudinal stability of the aircraft.

[0117] When the aircraft needs to perform pitch control, the lift thrusters of the forewing 2 work together with the tail thrusters to generate pitch torque. For example, if the aircraft needs to be controlled to pitch up, the lift thrusters adjust their speed to increase lift, while the tail tilt thrusters decrease lift, generating pitch torque to control the aircraft to pitch up. At the same time, the tail motors adjust their yaw to stabilize the aircraft's heading, and the lift thrusters adjust to stabilize the aircraft's roll angle.

[0118] When the aircraft needs to perform roll control, the lift thrusters of the front wing 2 are used to generate roll torque, and the tail wing is used to maintain directional stability. For example, if the aircraft needs to be controlled to roll to the right, the lift of the right lift thruster decreases and the left lift increases, generating a right roll torque; the tail motor adjusts the speed and yaw angle to keep the aircraft's pitch and yaw directions stable.

[0119] For the control logic of the control surfaces, similar to most eVTOL aircraft, the cruise mode of the aircraft in this study uses aerodynamic control surface deflection to generate control torque. To enhance the aircraft's control performance and better adapt to the maneuverability requirements of urban air transport missions, the aircraft adopts a hybrid control method to achieve attitude control.

[0120] When controlling the pitch of an aircraft, the front wing control surface 2 and the rear wing control surface 3 work together to deflect up and down to generate corresponding pitch forces. For example, when controlling the pitch of an aircraft, the front wing control surface 2 deflects down to increase the lift of the front wing, while the rear wing control surface 3 deflects up to reduce the lift of the rear wing. The two sets of control surfaces work together to generate a strong pitch control torque.

[0121] When controlling the aircraft to roll, the three control surfaces of the front and rear wings of the aircraft are differentially deflected to generate the corresponding rolling torque. For example, when controlling the aircraft to roll to the left, the three control surfaces of the front and rear wings of the aircraft are differentially deflected, the left control surface deflects upward and the right control surface deflects downward. The two sets of control surfaces work together to generate a strong rolling control torque.

[0122] When it is necessary to control the pitch and roll of an aircraft simultaneously, the effect of its control surface deflection is a vector superposition of the two situations mentioned above, which will not be elaborated further here.

[0123] The basic structure and principles of this scheme are as described above. The following section will provide the various indicators of this scheme to help you understand it in more detail.

[0124] First, the overall performance indicators are as follows, as can be found below.

[0125] Based on mission analysis and considering the characteristics of 10kg-class urban air traffic logistics transportation missions, the basic requirements for the overall performance of the aircraft are summarized below.

[0126] In terms of size and weight, this study defines the takeoff weight range and maximum size profile of the 10kg class logistics transport aircraft, which is most likely to be the first to be applied to missions, in order to meet the mission requirements of flexible takeoff and landing in small, simple takeoff and landing sites.

[0127] Regarding payload capacity, this study sets the effective payload mass and volume of the aircraft as the overall design indicators, considering the common payload types and volumes in urban logistics transportation tasks. Examining the current situation of single and multiple pieces of goods in urban logistics, it is believed that aircraft with a total cargo transport volume of 2-5 kg ​​and a maximum cargo hold size of approximately 0.4 m have high practical value.

[0128] In terms of cruise performance, a comprehensive evaluation of similar aircraft was conducted. Based on the length of urban transport routes and the distance, design parameters such as flight time and cruise speed were set. Furthermore, minimum flight speed and minimum feasible overload parameters were set for the aircraft to ensure that the aircraft has good low-speed performance and maneuverability, and better adapts to the mission requirements of shuttle take-off and landing between buildings.

[0129] In terms of maintainability, the aircraft is required to have a high degree of modular design and ensure that major components can be quickly disassembled and assembled; on the one hand, this is to reduce the maintenance cost of the aircraft, and on the other hand, to ensure that the aircraft has a low failure rate and a high uptime rate, thereby reducing the overall operating cost of the aircraft.

[0130] The overall performance indicators are as follows:

[0131] The aircraft's takeoff weight is required to be 9–15 kg, and its wingspan is required to be 3 m, to meet the basic characteristics of a 10 kg class aircraft, i.e., to meet the requirements for takeoff and landing in small spaces; the aircraft's effective payload is required to be 2–5 kg, and the cargo hold size is 0.4 m * 0.3 m * 0.08 m, to ensure a large cargo hold and large effective payload, meeting the requirements for diversified material delivery; the payload flight time is required to be 1–3 hours, and the cruising speed is required to be 14–18 m / s, to adapt to medium- and low-speed long-endurance flight, meeting the application of long-duration, multiple delivery missions; the minimum speed is required to be at least greater than 12 m / s, and the feasible overload is required to be at least greater than 2.75 G, to ensure low-speed performance and maneuverability, and to adapt to shuttle takeoff and landing between buildings in urban areas; a modular design is adopted, and the fuselage, wings and their transition sections can be replaced individually, to achieve low-cost, simple and quick maintenance, and reduce overall operating costs.

[0132] Secondly, regarding the aerodynamic parameters of this scheme, as shown below, they mainly include requirements for geometric dimensions, wing and fuselage geometric thickness, lift-drag characteristics, and handling characteristics.

[0133] The aircraft has a wingspan of 3m, a length of less than 1.5m, and a wing chord length of 300mm to meet the needs of small-space flight, transportation, and storage; the relative thickness of the fuselage and the relative thickness of the wings are at least 11% to meet the needs of heavy-load missions; the rated lift-to-drag ratio is greater than 12 to ensure long-endurance performance; the cruise speed is 14-18m / s, and the minimum flight speed is 12m / s to adapt to the needs of medium- and low-speed, flexible and maneuverable material transport missions; the pitch static stability is 5%-10%, and the yaw moment slope is greater than 0.005 to ensure stable flight.

[0134] Thirdly, there are the main structural design specifications for the aircraft, as shown below. These include constraints on the overall structural mass, requirements for structural load-bearing capacity, and other requirements to improve the maintainability of the aircraft.

[0135] The overall structural weight is 3.5kg to ensure load capacity; the cruise overload is 2.75G with a cruise overload deflection of less than 5cm to ensure cruise and maneuverability; the hover overload is 2.75G with a hover overload deflection of less than 1cm and a hover overload angular displacement of less than 1° to ensure VTOL flight stability and maneuverability; large component connections are detachable and can be disassembled and reassembled multiple times without damage, and reserved locations for load equipment, such as floor, cable trays, and positioning holes, are provided for convenient experimentation and application.

[0136] Finally, regarding the main mass distribution design parameters of the aircraft, see below.

[0137] Based on previous aircraft design experience, the mass allocation model roughly estimates the specifications and weight of each system and equipment of the aircraft. The detailed design will refer to the mass allocation model to ultimately achieve an aircraft design that meets the mass requirements, thus forming a design closed loop.

[0138] The rated takeoff weight of the aircraft is set, and the rated mass of the aircraft is also specified.

[0139] G = m e g

[0140] The rated mass of an aircraft is allocated as follows:

[0141] m e =m 动力 +m 结构 +m 设备 +m 载荷

[0142] The power system mainly consists of three components: batteries, motors, and motor drivers. Based on the typical weight distribution of batteries in eVTOL aircraft, the total battery weight range is determined. High-energy-density 21700 lithium iron phosphate batteries are selected, connected in parallel four times (12 sets in series) to obtain a high-energy battery pack with an output voltage of 44V, a maximum discharge current of 120A, and a total capacity of 20000mAh. The estimated rated takeoff weight of the aircraft is 12kg; therefore, two lift propulsion motors with a rated thrust of 3.5kgF and one external rotor brushless motor with a rated thrust of 3kgF are selected. The total mass of the motors is estimated based on the general parameters of these motors. Similarly, according to the motor specifications, the specifications of similar drivers are referenced, and the total mass of the drivers is estimated.

[0143] The power system consists of a 12S4P lithium battery weighing 3.7 kg, three motors for the propulsion system (two disc motors for the lift propulsion components weighing 3.5 kg each, and an external rotating motor for the tail propulsion component weighing 3 kg), and three 12S60A drives weighing 0.24 kg each, for a total power system mass of 4.84 kg. The structural components weigh 3.5 kg, the mechanism for securing the tail propulsion component weighs 0.2 kg, for a total mass of 3.7 kg. The avionics include a brushless tilt servo weighing 0.16 kg, a brushless yaw servo weighing 0.06 kg, a brushless control surface servo weighing 0.2 kg, a flight controller and various sensors weighing 0.5 kg, a 12V 6A regulated DC power supply weighing 0.1 kg, for a total avionics mass of 1.02 kg, a rated payload of 2 kg, and a rated takeoff weight of 12 kg.

[0144] The above describes the specific structure, principle, and design specifications of this scheme. The following section will use specific experimental verification to demonstrate the feasibility and beneficial effects of this scheme.

[0145] First, such as Figure 12 As shown, this scheme has been verified by Fluent calculations. Compared with traditional layout aircraft and tandem lifting body design aircraft, the compound tandem lifting body aircraft with refined aerodynamic design optimization has a higher lift-to-drag ratio, lift line slope, further reduced stall speed, and further increased stall angle of attack.

[0146] Secondly, the hovering stability of the aircraft was verified through the turntable experiment. The yaw control response was rapid and the yaw attitude angle followed closely, which verified that the use of a two-degree-of-freedom tail tilt thruster to reduce yaw control can significantly improve the hovering mode control performance of the aircraft.

[0147] Finally, rotorcraft mode flight experiments, mode transition flight experiments, and fixed-wing mode flight experiments were conducted respectively. The results verified that... Figure 13 As shown, its cruise power consumption is close to the theoretical minimum power consumption obtained by CFD calculation, but there is still some loss. This is due to the parasitic drag caused by the landing gear, communication equipment and other devices installed on the surface of the aircraft. The power consumption of the aerodynamic test aircraft during cruise is small, far lower than the maximum cruise power consumption required to meet the overall design mission requirements. This is due to the excellent cruise efficiency of the tandem-lift body aerodynamic configuration and the efficient aerodynamic design of this study.

[0148] The above description represents the preferred embodiments of the present invention. It should be noted that those skilled in the art can make various improvements and modifications without departing from the principles of the present invention, and these improvements and modifications are also considered to be within the scope of protection of the present invention.

Claims

1. A vertical takeoff and landing aircraft, characterized in that, This includes a blended wing-body lifting body fuselage, a front wing and a rear wing, a stabilizer tail, and a ring wing; The top of the lifting body fuselage protrudes upward to form an upper arc convex surface and an upper arc concave surface. The upper arc convex surface and the upper arc concave surface are arranged sequentially from the nose to the tail. The curvature of the upper arc convex surface is greater than that of the upper arc concave surface. The sides of the lifting body fuselage bulge outward to form a nose protrusion and a tail protrusion; The surface of the nose protrusion is rounded to the top surface of the lifting body fuselage and the front end surface of the lifting body fuselage. The forewing is mounted at the highest point of the nose protrusion, and the forewing is provided with a lift propulsion component. The nose protrusion narrows from its outer contour line to the highest point of the nose protrusion, and at least a portion of the nose protrusion extends to the tail protrusion. The extended portion of the nose protrusion narrows from the nose to the tail. The nose protrusion includes a smoothly transitioned mounting head section and an extended tail section; the forewing is mounted at the junction of the upper and lower parts of the mounting head section; the upper surface of the mounting head section smoothly transitions to the surface of the forewing, and the lower surface of the mounting head section connects to the surface of the tail protrusion at an obtuse angle, with the difference in curvature between the upper and lower parts of the mounting head section being within 2°; in the direction adjacent to the tail protrusion, the end of the extended tail section tapers into a pointed shape; the upper surface of the extended tail section smoothly transitions to the surface of the tail protrusion, and the lower surface of the extended tail section connects to the surface of the tail protrusion at an obtuse angle, with the upper curvature of the extended tail section being greater than the lower curvature of the extended tail section. The surface of the tail protrusion transitions to the top surface of the lifting body fuselage with an arc. The rear wing is installed at the highest point of the tail protrusion. The tail protrusion narrows from its outer contour line to the highest point of the tail protrusion. The lifting body fuselage is equipped with two stabilizer wings at its tail, the two stabilizer wings are arranged opposite each other, the wingtips of the two stabilizer wings are pointing upwards, and a tail propulsion component is installed between the two stabilizer wings. The ring wing is mounted on the bottom of the front wing, the root of the ring wing is connected to the fuselage, the wingtip of the ring wing is bent towards the top of the vertical take-off and landing aircraft, the wingtip of the ring wing is connected to the front wing, and a lift propulsion unit is mounted on the bottom of the ring wing; the wingspan of the ring wing is less than or equal to the wingspan of the front wing. The tail propulsion component is located below the tail of the lifting body fuselage; The tail propulsion component can tilt in two degrees of freedom toward the bottom of the fuselage and toward the side of the fuselage.

2. The vertical takeoff and landing aircraft according to claim 1, characterized in that, In the vertical direction, the mounting height of the front wing is higher than that of the rear wing. The mounting angles of the front wing and the rear wing are oriented in the same direction. Both the mounting angles of the front wing and the rear wing are less than 90 degrees, and there is an angle difference between the mounting angles of the front wing and the rear wing.

3. The vertical takeoff and landing aircraft according to claim 2, characterized in that, The front wing is mounted on the upper part of the nose protrusion, and the rear wing is mounted on the lower part of the tail protrusion; the height difference between the center of gravity of the front wing and the center of gravity of the rear wing is 30-65mm, and the angle difference between the mounting angle of the front wing and the mounting angle of the rear wing is 1-2°.

4. The vertical takeoff and landing aircraft according to claim 2, characterized in that, The tail section includes a smoothly transitioning middle section and an extended end section. In the direction adjacent to the nose section, the beginning of the middle section is an enlarged opening, and the pointed, narrowed end section is located on the V-shaped opening of the middle section. The middle section and the extended end section are matched and connected. The upper surface of the middle section smoothly transitions to the surface of the nose section, and the lower part of the middle section extends to the lower part of the nose section. The lower surface of the middle section forms an obtuse angle with the lower surface of the nose section, and the upper curvature of the middle section is greater than the lower curvature. The rear wing is mounted on the lower part of the extended end section, and the upper surface of the extended end section smoothly transitions to the surface of the rear wing.

5. The vertical takeoff and landing aircraft according to claim 1, characterized in that, The front wing has a front wingtip at its wingtip, and the rear wing has a rear wingtip at its wingtip; the wingtip of the front wingtip faces the bottom of the vertical take-off and landing aircraft, and the wingtip of the rear wingtip faces the top of the vertical take-off and landing aircraft.

6. The vertical takeoff and landing aircraft according to claim 1, characterized in that, The relative thickness of the lifting body fuselage is greater than 11%, and the length of the lifting body fuselage is less than 1.5m.

7. The vertical takeoff and landing aircraft according to claim 1, characterized in that, The relative thickness of the front wing and the rear wing is greater than 11%, the wing chord length of the front wing and the rear wing is 200-400 mm, and the wingspan of the front wing and the rear wing is 2-4 m.

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